This disclosure relates to a gas turbine engine. More particularly, the disclosure relates to a tip cooling configuration for an airfoil.
Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.
Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. The turbine vanes, which generally do not rotate, guide the airflow and prepare it for the next set of blades.
Turbine blades typically include internal cooling passages. Film cooling holes communicate cooling fluid from the cooling passages to areas on the exterior surface of the turbine blade that may experience undesirably high temperatures.
One high temperature area is the tip of the airfoil. A proposed cooling configuration uses a notch circumscribed about the entire perimeter of the airfoil at the tip, which permits leakage of the working fluid past the tip.